GB2596139A

GB2596139A - Fan blade tip operating clearance optimisation

Info

Publication number: GB2596139A
Application number: GB2009383.7A
Authority: GB
Inventors: Merriman Nicholas
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2020-06-19
Filing date: 2020-06-19
Publication date: 2021-12-22
Also published as: GB202009383D0

Abstract

A method of calculating a profile for the internal wall of a housing surrounding a turbine blade such that the blade tip clearance is optimised for a particular operating point, the method including selecting an operating point, determining an initial housing wall profile, determining an operating point casing profile, determining a blade tip clearance for the operating point, determining a profile correction, applying the profile correction to the initial casing profile to determine a modified casing profile that will provide an optimised blade tip clearance. There may also be a method of manufacturing a modified clearance profile, a Gas Turbine Engine fitted with the housing having the modified profile and a computer program to implement the method.

Description

FAN BLADE TIP OPERATING CLEARANCE OPTIMISATION

Field of the Disclosure

The present disclosure relates to a method of optimising a blade tip operating clearance between the blade tips of a fan of a gas turbine engine for an aircraft and the internal wall of a housing surrounding the fan. The present disclosure also relates to a method of manufacturing a gas turbine engine that incorporates the optimisation method and a resulting gas turbine engine. The present disclosure also relates to a computer program adapted to carry out the optimisation method and a computer readable medium having the computer program stored thereon.

Background of the Disclosure

An aircraft engine such as a gas turbine engine has a fan that rotates within a housing formed by a nacelle or fan case. The fan comprises a plurality of fan blades. A blade tip clearance is defined between the radial tips of the fan blades and the internal wall of the housing. It is desirable to minimise the blade tip clearance when the engine is in use. This helps to reduce leakage of air between the blade tips and the housing, thus improving fan efficiency. Reducing the blade tip clearance also reduces the noise produced by the engine.

While it is advantageous to reduce the blade tip clearance, use of small blade tip clearance causes the blade tips to come close to and sometimes contact the internal 25 wall of the housing. This may cause damage to the fan blades.

A general problem to be addressed is that of optimising the blade tip clearance so that detrimental damage to the fan blade is avoided by contact with the nacelle, while still providing efficient fan operation and low levels of noise.

Known solutions include the use of an attrition liner forming the inner surface of the housing that faces the blade tips. The blade tips are intended to pass as close as possible to the attrition liner when rotating. The attrition liner is designed to be abraded away by the fan blade tips during use of the engine. Rubbing between the blade tips and attrition liner changes the shape of the attrition liner. This allows a seal between the fan blades and the fan track liner to be formed and so improves the effectiveness of the fan in driving air through the engine.

Rubbing between the blade tips and housing to cause abrasion using this method is however undesirable as damage may be caused to the blade tips. It may also not be suitable for all types of fan blade construction or materials, some of which may be more susceptible to damage caused by contact with the housing.

It is an aim of the present disclosure to provide an improved method of optimising the fan blade tip clearance.

Summary of the Disclosure

According to a first aspect of the present disclosure, there is provided a method of optimising a blade tip operating clearance between the blade tips of a fan of a gas turbine engine for an aircraft and the internal wall of a housing surrounding the fan, the method comprising the steps of: a) selecting at least one engine operating point of the engine; b) determining an operating casing profile of the internal wall for the at least one selected operating point based on an initial casing profile of the housing; c) determining an operating fan blade tip profile for the at least one engine operating point; d) determining a casing profile radial correction based on the operating casing profile and the operating fan blade tip profile; e) applying the casing profile radial correction to the initial casing profile to determine a modified casing profile providing an optimised blade tip clearance.

By determining the modified casing profile in this way material can be added where there is otherwise a running clearance between the fan blade tips and inner surface of the housing at the selected operating point(s). Conversely, material can be removed where there is otherwise rubbing between the fan blade tips and inner surface of the housing. This means that when the inner surface of the housing having the modified casing profile changes shape during operation an optimised blade tip spacing is still provided. This allows an optimal blade tip spacing to be provided, without relying on rubbing between the blade tips and inner surface of the housing to create the optimal shape of the inner surface by abrasion during use of the engine.

Step a) may comprise selecting a plurality of engine operating points. Step b) may comprise determining the operating casing profile over all of the plurality of engine operating points. Step c) may comprise determining the operating fan blade tip profile over all of the plurality of engine operating points.

The operating casing profile may be determined by determining the minimum radius of the casing profile for each point around the rotational axis of the engine occurring over all of the plurality of engine operating points. The operating fan blade tip profile may be determined by determining the maximum radius of the fan blades of the fan occurring over all of the plurality of engine operating points.

The plurality of engine operating points may include operating points that represent an entire flight cycle of the gas turbine engine.

The operating casing profile may be determined based on the effects of the engine weight and/or the thrust provided by the engine at the, or each, operating point.

The operation casing profile may be determined based on rapid transient effects. The rapid transient effects may include wind gust loading.

The method may further comprise repeating at least steps b), d) and e) for a plurality of iterations. The resulting modified casing profile for each iteration may be used as the initial casing profile for the following iteration.

The method may further comprise repeating at least steps b), d) and e) until a constant modified casing profile is determined.

The casing profile may define the shape of the internal surface of the housing as the radial distance of the inner surface from a rotational axis of the engine at associated angles of rotation around the rotational axis.

The shape of the internal surface of the housing may be defined by the casing profiles at a plurality of axial points at or between the axial positions of the leading edges and the trailing edges of the fan blade tips. The operating fan blade tip profile may define corresponding radii of the tips of the fan blades at or between the axial positions of the leading and trailing edges of the fan blade tips.

The modified casing profile may be non-axisymmetric about the rotational axis of the engine.

The modified casing profile may have a minor axis and a major axis. The minor axis and major axis may be perpendicular to one another. The major axis may be greater in length than the minor axis. The major and minor axes may extend across the modified casing profile and intersect (and be orthogonal to) the rotational axis of the engine.

The gas turbine engine may have a vertical axis. The vertical axis is aligned with vertical when the gas turbine engine is mounted to an aircraft at rest, and intersects the rotational axis of the engine. The minor axis of the modified casing profile may be orientated in a 90 degree range measured clockwise from the vertical axis of the engine when viewed in a rearward direction from a position forward of the fan.

The modified casing profile may have a generally elliptical shape.

The modified casing profile may vary in radius from axisymmetric by 15 mm or less at any point around the rotational axis of the engine.

The fan may comprise a plurality of fan blades. Each of the fan blades may be formed at least in part from a composite material. Each fan blade may comprise a body formed from a composite material and a leading edge formed from a metallic material.

The fan may have a diameter in the range of between 240 cm and 380 cm. The fan 25 may have a diameter between 240 cm and 280 cm or between 330 cm and 380 cm.

The gas turbine engine may comprise an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The fan may be located upstream of the engine core. The fan may comprise a plurality of fan blades.

The gas turbine engine may further comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft.

According to a second aspect of the present disclosure, there is provided a method of manufacturing a gas turbine engine for an aircraft, the gas turbine engine comprising a fan having a plurality of fan blades, and a housing surrounding the fan, the housing having an internal surface facing the blades of the fan, the method comprising: a) obtaining a modified casing profile associated with the gas turbine engine, the modified casing profile being determined using the method of the first aspect or 10 any of the statements above; and b) manufacturing the housing for the fan of the gas turbine engine, the housing having an internal wall shaped according to the modified casing profile.

According to third aspect of the present disclosure, there is provided a gas turbine engine for an aircraft, the gas turbine engine manufactured according to the method of the second aspect.

According to a fourth aspect of the present disclosure, there is provided a computer program having instructions adapted to carry out the method according to the first aspect or any of the above statements.

According to a fifth aspect of the present disclosure, there is provided a computer 20 readable medium, having a computer program recorded thereon, wherein the computer program is adapted to make the computer execute the method of the first aspect or any of the above statements.

According to a sixth aspect of the present disclosure, there is provided a gas turbine engine for an aircraft, the gas turbine engine comprising: a fan having a plurality of fan blades; and a housing surrounding the fan, the housing having an internal surface facing the blades of the fan, wherein the internal surface of the housing has a nonaxisymmetric profile about a rotational axis of the engine.

The non-axisymmetric profile may be present where the gas turbine engine is at rest (e.g. not operating), and/or before any abrasion of the internal surface has occurred during operation of the engine.

The non-axisymmetric profile may have a minor axis and a major axis. The minor axis and major axis may be perpendicular to one another. The major axis may be greater in length than the minor axis. The gas turbine engine may have a vertical axis. The minor axis of the modified casing profile may be orientated in a 90 degree range measured clockwise from the vertical axis of the engine when viewed in a rearward direction from a position forward of the fan.

The non-axisym metric profile may be generally elliptical in shape.

The non-axisymmetric profile may vary in radius from axisymmetric by 15 mm or less at any point around the rotational axis of the engine.

As noted elsewhere herein, the present disclosure may relate to a gas turbine engine.

Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a "planetary" or "star" gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a "star" gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The corn bustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes On that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage 15 may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U,2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Wu, is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct 113 may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest-pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg-ls, 105 Nkg-ls, 100 Nkg-ls, 95 Nkg-ls, 90 Nkg-ls, 85 Nkg-ls or 80 Nkg-ls. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg-ls to 100 Nkg-ls, or 85 Nkg-ls to 95 Nkg-ls. Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330kN to 420 kN, for example 350kN to 400kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C (ambient pressure 101.3kPa, temperature 30 degrees C), with the engine static.

In use, the temperature of the flow at the entry to the high-pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on zo the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example, at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium-based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel-based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades.

As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the "economic mission") of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint -in terms of time and/or distance-between top of climb and start of descent. Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example, where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide -in combination with any other engines on the aircraft -steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000m to 15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30kN to 35kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000ft (11582m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50kN to 65kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000ft (10668m).

In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

According to an aspect of the present disclosure, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein.

According to an aspect of the present disclosure, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number).

According to an aspect of the present disclosure, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Brief Description of the Drawings

Embodiments will now be described by way of example only, with reference to the Figures, in which: Figure 1 is a sectional side view of a gas turbine engine; Figure 2 is a close-up sectional side view of an upstream portion of a gas turbine 20 engine; Figure 3 is a partially cut-away view of a gearbox for a gas turbine engine; Figure 4 shows a close-up section view of the blade tip region of a fan blade; Figure 5 shows a method of optimising a fan blade tip operating clearance; Figure 6 is a cross section through the fan and nacelle inner surface of the gas turbine engine of Figure 1 in a plane normal to the rotational axis of the engine illustrating an initial non-operating casing profile and non-operating fan tip profile; Figure 7 is another a cross section through the fan and nacelle inner surface of the gas turbine engine of Figure 1 in a plane normal to the rotational axis of the engine illustrating a casing profile and fan tip profile at a take-off operating point of the engine; Figure 8 shows a casing profile radial correction plotted against rotation around the rotational axis of the engine; Figure 9 shows a cross section corresponding to that of Figure 6 illustrating a modified casing profile determined according to the method of Figure 5; Figure 10 shows another cross section corresponding to that of Figure 6, illustrating the modified casing profile at the take-off operating point demonstrating blade tip operating clearance optimisation; Figure 11 shows a modified casing profile at the non-operating condition, a modified casing profile at the selected operating point, an operating fan blade tip profile, an operating casing profile, a non-operating fan blade tip profile and a non-optimised casing profile each as a plot against rotation around the rotational axis of the engine; Figure 12 shows a cross section corresponding to Figure 6 illustrating the casing profile and fan blade tip profile at a climb phase operating point of the engine; zo Figure 13 shows another method of optimising a fan blade tip operating clearance; Figure 14 shows a flight cycle of a gas turbine engine made up of phases having associated engine operating points; Figure 15 shows a cross section corresponding to that of Figure 6 illustrating a modified casing profile determined according to the method of Figure 13; Figures 16 and 17 show further examples of an operating casing profile and a modified casing profile determined using methods of the present application; and Figure 18 shows a method of manufacturing a gas turbine engine

Detailed Description

Figure 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11 comprises, in axial flow series, a low-pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low-pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low-pressure turbine 19 via a shaft 26 and an epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the low-pressure compressor 14 and directed into the high-pressure compressor 15 where further compression takes place. The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low-pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust.

The high-pressure turbine 17 drives the high-pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Figure 2. The low-pressure turbine 19 (see Figure 1) drives the shaft 26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclic gear arrangement 30. Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

Note that the terms "low-pressure turbine" and "low-pressure compressor' as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the "low-pressure turbine" and "low-pressure compressor" referred to herein may alternatively be known as the "intermediate pressure turbine" and "intermediate pressure compressor". Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail in Figure 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in Figure 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox 30 generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in Figures 2 and 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in Figures 2 and 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10.

By way of further example, the connections (such as the linkages 36, 40 in the Figure 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Figure 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement 113 of output and support linkages and bearing locations would typically be different to that shown by way of example in Figure 2.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Figure 1 has a split flow nozzle 18, 20 meaning that the flow through the bypass duct 22 has its own nozzle 18 that is separate to and radially outside the core engine nozzle 20. However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine. In some arrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in Figure 1), and a circumferential direction (perpendicular to the page in the Figure 1 view). The axial, radial and circumferential directions are mutually perpendicular.

The fan 23 comprises a plurality of fan blades 41. The fan blades 41 are attached to a central hub or disc from which they extend in a radial direction. The fan 23 is arranged to rotate about the rotational axis 9 of the engine. As shown in Figures 1 and 2, the nacelle 21 surrounds the fan 23, with an inner wall 8 of the nacelle 21 facing the radial tips of the fan blades.

Figure 4 illustrates a section view of a fan blade 41 showing the blade tip region in more detail. The fan blade 41 comprises a tip surface 42 that extends between a leading edge 43 and a trailing edge 44 of the fan blade 41. The tip surface 42 also extends across the thickness of the fan blade 41 between the suction side and pressure side. A blade tip clearance D is defined as the radial distance between a point on the blade tip surface 42 and a corresponding point on the inner surface 8 of the nacelle 21.

As shown in Figures 1 and 4, the internal wall 8 of the nacelle 21 comprises an attrition liner 7. The attrition liner 7 is formed of a material that is robust to being contacted by the fan blades 41, and forms part of the inner surface of the nacelle that may come into contact with the blade tip surface. The attrition liner 7 can therefore protect the rest of the nacelle 21 from contact with the tips of the fan blades. The attrition liner 7 may be formed of a material that is abradable by contact with tip surface 42 of the fan blades as is known in the art. The blade tip clearance D may therefore be the radial distance between the blade tip surface 42 and the inner surface of the attrition liner 7 forming the associated part of the inner surface 8 of the nacelle 21.

The part of the nacelle 21 that surrounds the fan is considered to form a housing for the fan. In some embodiments, the gas turbine engine may comprise a fan case which surrounds the fan. The fan case is in turn mounted to or housed by the nacelle. The inner surface 8 facing the blade tip surfaces 42 may therefore be the inner surface of any housing that surrounds the fan, for example either the inner surface of a nacelle, fan case or other structure.

The blades of the fan 23 are manufactured at least in part from a composite material as described elsewhere herein. The composite material may be a metal matrix composite and/or an organic matrix composite, such as carbon fibre. In the presently described embodiment, the fan blades comprise two regions manufactured using different materials. The fan blade has a protective leading edge, that is manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. The method of optimising the blade tip clearance described may be used in combination with the use such composite fan blades because they are more susceptible to damage caused by rubbing against the inner surface of the nacelle. Where a protective leading edge is provided on the fan blade this may not extend to the tip surface of the fan blade, and may specifically be susceptible to damage caused by rubbing. The method of the present application is not limited to only this type of fan blade, and provides improved fan tip optimisation for all types of fan blades formed from any suitable material.

zo The present application relates to a method of optimising the blade tip clearance D. Figure 5 illustrates a method 100 of optimising the blade tip clearance D. In order to optimise the blade tip clearance during operation of the engine, a modified profile, referred to as a modified casing profile, of the inner surface 8 of the nacelle 21 facing the blade tip surface is determined. The casing profile defines the shape of the inner surface 8 of the nacelle 21. In the described embodiment, the casing profile defines the shape of the inner surface of the attrition liner 7 that forms the part of the inner surface 8 of the nacelle facing the tip surfaces 42 of the fan blades 41. The shape of the inner surface 8 of the nacelle 21 is defined as the radial distance of the inner surface 8 from the rotational axis 9 of the engine 10 at respective angles of rotation around the rotational axis 9.

Before optimisation, the inner surface 8 is axisymmetric around the rotational axis 9. An example of an initial, non-optimised casing profile 46 with the engine in a non-operating rest condition is shown in Figure 6. Figure 6 illustrates a schematic cross section through the fan 23 and inner surface 8 of the nacelle 21 in a plane normal to the rotational axis 9 of the engine. As can be seen in Figure 6, the casing profile is cylindrical about the rotational axis 9 of the engine and has a constant blade tip clearance D at all points around the rotational axis. The casing profile may not be cylindrical, and may vary in radius along the axial length of the engine as shown in Figure 4. The non-optimised casing profile is defined by the radial distance RCN at respective angles around the rotational axis 9. The non-optimised casing profile shown in Figure 6 corresponds to the casing profile before any change in shape (e.g. by abrasion of the attrition liner) has occurred during use of the engine.

Figure 6 also illustrates a fan blade tip profile 48. The blade tip profile 48 is defined by the circle swept out by a respective point on the tip surface 42 of the fan blades. The blade tip profile 48 in the non-operating rest condition is illustrated in Figure 6 as having a constant radius RFN around the central rotational axis 9.

Method 100 comprises the step of selecting 102 an operating point of the engine. The engine operating point refers to different operating conditions of the engine, and may include operating the engine at idle, taxi, take-off, climb, cruise, decent and landing phases as will be described later (and may be referred to as an engine operating condition). The operating points described herein are provided as examples only, with others being possible depending on the implementation and intended use of the engine.

The method 100 further comprises determining 104 an operating casing profile 50 of the inner surface 8 of the nacelle 21. The operating casing profile 50 is determined at the operating point selected in step 102, and defines the shape of the nacelle inner surface 8 when the engine is operating under the conditions of the selected operating point. An example of a casing profile 50 at a take-off operating point is shown in Figure 7. As can be seen in Figure 7, during operation of the engine the shape of the nacelle 21 and its inner surface 8 facing the blade tips of the fan changes so that it is no longer axisymmetric about the rotational axis 9 of the engine. In other words, the radius RCT (i.e. the radius at the take-off operating condition) of the inner surface 8 of the nacelle varies between different angles of rotation around the rotational axis 9 of the engine.

The change in shape between the non-operating point and at the selected operating point can be expressed as a change in radius of the casing, ORC(0), given by: ORC(0) = RC+(9) -RCN(0) (1) for respective angles around the rotational axis 9 of the engine.

The method 100 further comprises determining 106 an operating fan blade tip profile 52 at the engine operating point selected in step 102. The operating fan blade tip profile 52 is illustrated in Figure 7, and is defined as the radius RFT at respective values of the angle around the engine rotational axis 9. As can be seen in Figure 7, the fan operating blade profile 52 has a greater radius compared to the non-operating fan tip profile because of expansion of the fan blades and movement of the fan during rotation. The fan blade tip profile and casing profile are in the same plane of the engine.

Following step 106, the method 100 comprises a step of determining 108 a casing profile radial correction 54 based on the operating casing profile 50 and the operating fan blade tip profile 52. The radial correction is determined by finding the difference between the operating casing profile 50 and the operating fan profile 52 at each point around the rotational axis 9 (i.e. RF+(e) -RC+(e)). The radial correction corresponding to Figure 7 is illustrated in Figure 8, which shows the radial correction 54 on a plot of the change in radius, 6R, where R is the radius from the rotational axis 9 and angle, 0, of rotation around the rotational axis 9. An angle of zero or 27c corresponds to a vertical axis of the engine when it is mounted to an aircraft wing.

Once the radial correction has been determined, the method 100 comprises determining 110 a modified casing profile 56 that provides an optimised blade tip clearance. The modified casing profile 56 is illustrated in Figure 9. The modified casing profile 56 is determined based on the casing profile radial correction 54. The modified casing profile can be found by applying the radial correction 54 to the non-optimised casing profile 46. The modified casing profile RCM(0) can be found using the following expression RCM(0) = RCN(0) + (RFT(0) -RC-F(0)) + 6 (2) The term 5 represents a predefined blade tip spacing at the selected operating point.

By modifying the non-optimised casing profile in this way material is added where there is otherwise a running clearance between the fan blade tips and inner surface of the nacelle at the selected operating point. Conversely, material is removed where there is otherwise rubbing between the fan blade tips and inner surface of the nacelle. This means that when the inner surface of the nacelle having the modified casing profile changes shape during operation an optimised blade tip spacing is still provided.

113 The modified casing profile referred to herein is the shape of the housing before any abrasion that may occur during operating of the engine that causes a further change in shape of the inner surface of the housing.

An example of a nacelle having an inner surface shaped according to the modified casing profile is illustrated at the selected operating point in Figure 10. This figure illustrates the change in shape experienced by the nacelle 21 (and hence its inner surface 8) between the non-operating point of Figure 9 and the selected operating point. The modified casing profile is labelled 56 in Figure 9 and 56 in Figure 10 to reflect the change in shape. The modified casing profile 56 at the non-operating point, the modified casing profile at the selected operating point 56, the operating fan blade tip profile 52, the operating casing profile 50, the non-operating fan blade tip profile 48 and the non-optimised casing profile 46 are also illustrated in Figure 11, which shows a plot of radius R from the rotation axis 9 against rotation 0 around that axis.

The change in shape experienced by the nacelle having the modified casing profile can be considered to be the same as that experienced by the non-modified inner surface of the nacelle because the radial correction applied has no significant change in the overall behaviour of the nacelle (e.g. because only a small change in shape is applied and/or only the non-structural attrition liner 7 is modified).

At the selected operating point, the optimised casing profile RC0(0) can be expressed as: RCo(0) = RCM(0) + SRC(0) (3) By substituting expressions (1) and (2) above it can be seen that RCo is given by: RC0(9) = RCN(B) + (RFT(B) -RC1(0)) + 5 + ORC(e) = RCN(e) + RF+(e) -SRC(9) -RCN(8) + 6 + 5RC(e) = RFT(9) + 6 At the selected operating point, the modified casing profile RCo(9) differs from the fan tip profile RFT(e) by the predefined fan tip spacing 6. The method 100 therefore provides an optimisation of the fan tip spacing at the selected operating point that takes into account of the change in shape of the nacelle during operation of the engine.

If the value of 5 is set to zero, the modified casing profile is such that, at the selected operating point, the operating casing profile is the same as the operating fan tip profile.

The value of 5 can be adjusted to provide a desired blade tip spacing at the selected operating point. For example, a clearance between the blade tips and inner surface of the nacelle can be provided at the selected operating point by selecting a positive value for 6. Conversely, a degree of rubbing between the fan blade tips and inner surface of the nacelle may be provided if a negative value of 5 is chosen. This therefore allows the blade tip spacing at the selected operating point to be optimised.

In step 104 the operating casing profile 50 at the selected operating point is found by taking into account one or more of the following effects on the shape of the nacelle: i) Engine weight. The weight of the engine when it is attached to an aircraft wing will generate stresses within the nacelle causing sagging which will affect its shape.

ii) Thrust provided by the engine at the selected operating point. At different levels of thrust a different aerodynamic load will be applied to the nacelle. Varying levels of aerodynamic load will cause the nacelle to change shape between different operating points associated with different thrust levels.

iii) Rapid transient effects. These include effects that cause a short, temporary change in the shape of the nacelle. Rapid transient effects taken into account by the present method include wind gust loading. Wind gust loading causes a change in the incidence angle of air on the wings of an aircraft to which the engine is mounted. This causes varying levels of lift to be generated by the wings, and in turn causes changes in shape of the engine as those lift forces are transmitted to the engine via its mounting. Rapid transient effects can be modelled using Monte Carlo modelling techniques in which random changes in shape to the nacelle is introduced.

These effects on the shape of the nacelle can be modelled using stress and aeromechanical methods that are known in the art. The operating casing profile is determined based on the initial, non-optimized axisymmetric casing profile as a starting point to which one or more of the above effects are applied. The operating casing profile may be determined by taking into account any one or more of: the flight profile (e.g. which mode of flight or engine operating condition is experienced), how the engine is supported on the aircraft by its respective supporting pylon (e.g. the pylon can support the engine by the engine core only or by the core and fan case combined), and the intake design.

In one embodiment, the operating casing profile is determined based on the effects of engine weight and thrust, but not rapid transient effects. In such an embodiment, a small level of rubbing between blade tip surface and the inner surface of the nacelle will occur during rapid transient effects when the engine is in use. This allows the shape of the inner surface to be optimised by a small level of rubbing, rather than by taking into account rapid transient effects which, due to their varying nature, may be difficult to model as accurately as shape changes resulting from effects such as the engine weight and thrust level.

In other embodiments, transient effects are taken into account to mitigate reduction in performance over a number of flight cycles. This may be done by repeating the method 200 while introducing random transient affects. The method may be repeated iteratively by using the resulting modified casing profile as the initial casing profile for the following iteration. The method may be repeated over a number of iterations until a constant modified profile is found, or repeated over a set number of iterations that represent the use of the engine over its lifetime.

In step 106 of the optimising method 100 the operating fan tip profile is determined by taking into account changes to the area swept out by the tips of the fan blades during operation of the engine. These changes may be caused by radial expansion of the fan blades 41 resulting from centrifugal loading on the fan blades as they rotate, changes in temperature, and movement of the fan away from the rotational axis 9. The operating fan tip profile may be determined using suitable stress and aeromechanical modelling methods similarly to the operating casing profile.

In the method described above the change in shape of the inner surface of the nacelle at one selected operating point is taken into account. The selected operating point may be the cruise condition so that the blade tip clearance is optimised for the operating point at which the engine will spend the majority of time operating at. Other operating point may be chosen.

The change in shape of the nacelle may differ between different operating points. Figure 12 illustrates the change in shape at a climb phase operating point. As can be seen in Figure 12 the change in shape differs from that of the take-off phase operating point. At each of the operating conditions the operating casing profile is generally elliptical in shape, with different orientations of the elliptical shape around the rotational axis of the engine at different operating conditions.

In other embodiments, the method may take into account the change in shape of the inner surface 8 of the nacelle 21 and fan 23 at a plurality of different engine operating points. An example of such an embodiment 200 is illustrated in Figure 13. Method 200 includes steps that correspond to those of the method illustrated in Figure 5, and have been labelled with corresponding reference numbers accordingly. The features described above in connection with the method of Figure 5 may apply to that of Figure 13, and vice versa.

Referring to Figure 13, the method 200 comprises steps of selecting 202 a plurality of operating points, determining 204 an operating casing profile, determining 206 an operating fan blade tip profile, determining 206 a casing profile radial correction and determining a modified casing profile 208.

Step 202 comprises selecting a plurality of engine operating points over which the operating casing profile is to be determined. The operating points correspond to different engine operating points at various points of an operating cycle of the engine. A typical operating cycle 300 is illustrated in Figure 14, which shows a plot of aircraft altitude (A) against time (T). The operating cycle 300 shown in Figure 14 includes an at-rest, powered down condition 302, taxi-out phase 304, take-off phase 306, initial 113 climb phase 308, first cruise phase 310, mid-cruise climb phase 312, second cruise phase 314, first decent phase 316, hold phase 318, second decent phase 320, landing phase 322 and taxi-in phase 324. Each phase of the operating cycle has an associated engine operating point or operating conditions. In the present example, selecting a plurality of engine operating points comprises selecting all separate engine operating points that are experienced over the engine operating cycle. The flight cycle shown in Figure 14 is an example only, with others possible. A suitable flight cycle and operating points are chosen for the particular engine with which the method is being performed so that all operating points or conditions that will be experienced by the engine can be taken into account when determining the modified casing profile. In some embodiments however, not all of the entire operating cycle maybe covered. Engine operating points may therefore be selected for some, but not all, of the corresponding phase of the flight cycle. This may allow only the most significant operating cycles (e.g. cruise and landing) to be taken into account).

The method 200 further comprises determining 204 an operating casing profile 50 of the inner surface 8 of the nacelle 21. The operating casing profile 50 is determined over all of the operating points selected in step 202. In this embodiment, the operating casing profile 50 represents the minimum radius of the casing profile for each point around the rotational axis 9 that occurs over all of the operating points. A separate operating casing profile for each of the operating points is determined from the non-optimised casing profile 46 using appropriate stress and aeromechanical methods that take into account one or more of the effects on nacelle shape described above. An overall operating casing profile for all of the operating points can then be derived from the separate operating casing profiles.

The method 200 further comprises determining 206 an operating fan blade tip profile 52 over the engine operating points selected in step 202. In this embodiment, the operating fan blade tip profile 52 is the maximum radius of the fan blades that occurs over all of the operating points selected in step 202 and therefore represents the maximum extent of the range in variation of the fan blade tips.

An example of the operating fan blade tip profile 52 and the operating casing profile 50 resulting from step 204 and 206 is shown in Figure 15. In this example, the operating points selected are the take-off flight phase and climb flight phases. The broken lines in Figure 15 shows the operating casing profile for the take-off and climb operating points corresponding to those shown in Figures 7 and 12. As can be seen in Figure 15, the operating casing profile 50 over both of these operating points (shown in the solid line) is the smallest radius RCF which occurs at each point around the rotational axis 9.

Following steps 204 and 206, the method 100 comprises determining 208 a casing profile radial correction based on the operating casing profile 50 and the operating fan blade tip profile 52. This is determined in the same was as for the method of Figure 5 i.e. by subtracting the radius of the operating casing profile 50 from the radius of the operating fan profile 52 at each point around the rotational axis 9 of the engine.

Finally, once the radial correction has been determined, the method 200 comprises determining 210 a modified casing profile 56 that provides an optimised blade tip clearance similarly as described in connection with the method of figure 5.

The method 200 therefore calculates the maximum additional depth of liner material that can be added and no (or a predetermined small amount) of blade tip rubbing occur. Where the blade tips would otherwise rub, the minimum amount of material can be removed to avoid rubbing. The resulting modified casing profile will then result in the minimum cruise tip clearance possible and hence reduce fuel burn, with no requirement for the blade tips to rub to otherwise set the optimal shape by abrasion of the inner surface of the nacelle.

As discussed above, the modelling of transient effects on the change in shape of the nacelle may involve adding randomly occurring changes in shape caused by short duration effects such as wind gusts. In some embodiments, the method 100, 200 of determining a modified casing profile can be repeated over a number of iterations until a stable value of the modified casing profile is determined. In this embodiment, the method 100, 200 is repeated with the modified casing profile determined in one iteration used as the starting non-optimised casing profile for the next iteration. This may provide a more accurate determination of the modified casing profile that optimises the blade tip clearance over accumulated flight cycles. The method may be repeated over a number of iterations that represent the lifetime use of the engine.

In yet other embodiments, the method 100, 200 may be iterated over a number of different flight cycle operating points to model use of an engine for different types of flight. In yet other embodiments, the method 100, 200 may be iterated over different types of aircraft for the same type of engine or for a fleet of aircraft.

When iterating the method 100, 200 only steps 104, 108 and 210 or 204, 208, 210 may be repeated for a plurality of iterations. In other words, the same engine operating point(s) and operating fan profile may be used for each iteration In other 20 embodiments, all of the steps may be repeated as necessary.

In the embodiments described above, the fan tip profiles and casing profiles are determined in a plane normal to the rotational axis, for example in the plane marked X in Figure 4. This results in a modified casing profile to provide optimisation for the tip clearance D to the leading edge of the fan blade. In other embodiments, the method of determining the modified casing profile is performed at a plurality of axial planes at axial positions at or between the leading and trailing edges of the fan blade tip surface. This allows the blade tip spacing to the optimised along the length of the tip surface 42 of the fan blade 41. By repeating the method at different axial positions, the differing behaviour of the fan blade and variations in fan blade spacing along the fan tip surface can be provided. The shape of the internal surface 8 of the housing surrounding the fan is therefore defined by the casing profile at a plurality of axial points at or between the axial positions of the leading edges and the trailing edges of the fan blade tips.

The fan blade tip profile defines corresponding radii of the tip surfaces of the fan blades at or between the axial positions of the leading and trailing edges of the fan blades.

In the example shown in Figure 15, the change in shape between the operating casing profile for the take-off and climb operating points is exaggerated to aid illustration.

Figure 16 shows another view from the front of the engine 10 (i.e. from a position forward of the engine looking in a rearward direction of the engine towards the fan) to illustrate how the housing surrounding the fan changes in shape for different engine operating points. The initial casing profile 46 defined above is shown in Figure 16, with the fan absent for ease of illustration.

The operating casing profiles 50 determined according to the method described herein for each different engine operating condition have a generally elliptical profile, one of which is illustrated in Figure 16. The orientation of the operating casing profile(s) 50 is defined according to vertical and horizontal axes of the engine when mounted to an aircraft. The engine has a horizontal axis U and a vertical axis V corresponding to the orientation of the engine when mounted to an aircraft and the aircraft is at rest. Orientation angles defined herein are measured from the vertical axis in a clockwise rotation when looking in a rearward direction towards the fan from a position forward of the engine (i.e. alignment with the vertical axis corresponds to zero degrees).

The operating casing profiles 50 determined by the present method at different operating conditions vary in orientation about the rotational axis of the engine over a 90-degree range (labelled i3 in Figure 16). In other words, the orientation of the respective axis of the ellipses forming the operating casing profiles fall within a 90-degree range of angles about the rotational axis of the engine. For the engine illustrated in Figure 16, the minor axes (labelled 'a' in Figure 16) of the ellipses forming each operating casing profile lie within the 90-degree range measured clockwise from the vertical axis. For example, the operating casing profile for the take-off operating condition is orientated so that its minor axis is at an angle of 60-degrees measured clockwise from the vertical axis V as illustrated by the angle labelled a in Figure 16.

Figure 17 illustrates a modified casing profile 56 for the engine of Figure 16 determined using the method of the present application. In this example, the modified casing profile 56 is based on a plurality of individual elliptical operating casing profiles (at different operating conditions) falling within the 90-degree range described above in connection with Figure 16.

The modified casing profile 56 is non-axisymmetric about the rotational axis of the engine. It has a minor axis (labelled 'a') and a major axis (labelled 'b') that are perpendicular to each other, where the major axis is greater in length than the minor axis. In various implementations, the minor axis of the modified casing profile 56 resulting from the optimisation methods defined herein may fall within a 90-degree range of angles about the rotational axis of the aircraft. Specifically, the minor axis is orientated in a 90-degree range measured clockwise from the vertical axis V labelled y in Figure 17.

The modified casing profile 56 has a generally elliptical shape. As can be seen in Figure 17, it is more oblate or flatted in the direction of the minor axis compared to the individual elliptical operating casing profiles at each operating condition from which it is derived. Where the modified casing profile is defined over a single operating condition it may be an ellipse derived from the elliptical operating casing profile at that engine operating condition. By 'generally elliptical' in shape we therefore mean an elliptical or approximately elliptical shape including a flattened ellipse as shown in Figure 17.

As already described, the casing profile radial correction 54 varies around the rotational axis of the engine between a minimum radius and a maximum radius. In various implementations, the resulting modified casing profile 56 varies in radius from axisymmetric by 15 mm or less (i.e. greater than zero mm and less than or equal to 15 mm) at any point around the rotational axis of the engine. The difference between the length of the minor axis and the major axis of the modified casing profile 56 is therefore 30 mm or less. The casing profile radial correction 54 has a corresponding change in radius 6R that has as a peak-to-peak amplitude L of 15 mm or less as shown in Figure 8.

The ranges in the previous paragraph may be the case for a number of different engine and fan sizes, including engines having any of the fan diameters defined anywhere herein. In one example, the above range is suitable for an engine having a fan diameter in the range of 240 cm to 380 cm, specifically between 240 cm and 280 cm or between 330 cm and 380 cm.

It will be appreciated that the optimisation steps described herein may be used as part of a method of manufacturing a gas turbine engine. Such a method is illustrated in Figure 18 which illustrates a method 400 of manufacturing a gas turbine engine for an aircraft, the method comprising obtaining 402 a modified casing profile associated with the gas turbine engine and manufacturing 404 a fan housing (e.g. a fan case or nacelle) of the gas turbine engine. The modified casing profile is obtained 402 using the method of optimising the blade tip operating clearance defined anywhere herein. The fan housing is manufactured 404 having an internal wall shaped according to the modified casing profile using any suitable technique known in the art.

It will also be appreciated that any of the methods of optimising a blade tip operating clearance defined or claimed anywhere herein may be a computer implemented method of optimising a blade tip operating clearance. The optimisation methods of the present application may be encompassed in computer-implemented code and may be stored on a computer-readable medium. The method may be implemented on a general purpose computer system, e.g. comprising a processing unit, memory, user interface means such as a keyboard and/or mouse, and display means.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and sub-combinations of one or more features described herein.

Claims

CLAIMS1. A method (100, 200) of optimising a blade tip operating clearance (D) between the blade tips (42) of a fan (23) of a gas turbine engine (10) for an aircraft and the internal wall (8) of a housing (21) surrounding the fan (23), the method comprising the steps of: a) selecting (102) at least one engine operating point of the engine; b) determining (104) an operating casing profile (50) of the internal wall (8) for the at least one selected operating point based on an initial casing profile (46) of the housing (21); c) determining (106) an operating fan blade tip profile (52) for the at least one engine operating point; d) determining (108) a casing profile radial correction (54) based on the operating casing profile (50) and the operating fan blade tip profile (52); e) applying (110) the casing profile radial correction (54) to the initial casing profile (46) to determine a modified casing profile (56) providing an optimised blade tip clearance.
2. The method (100, 200) of Claim 1, wherein: step a) comprises selecting (202) a plurality of engine operating points; step b) comprises determining (204) the operating casing profile (50) over all of the plurality of engine operating points; and step c) comprises determining (206) the operating fan blade tip profile (52) over all of the plurality of engine operating points.
3. The method (100, 200) of Claim 2, wherein: the operating casing profile (50) is determined by determining the minimum radius of the casing profile for each point around the rotational axis of the engine occurring over all of the plurality of engine operating points; and the operating fan blade tip profile (52) is determined by determining the maximum radius of the fan blades (41) of the fan (23) occurring over all of the plurality of engine operating points.
4. The method (100, 200) of Claim 2 or Claim 3, wherein the plurality of engine operating points includes operating points that represent an entire flight cycle (300) of the gas turbine engine (10).
5. The method (100, 200) of any one of the preceding claims, wherein the operating casing profile (50) is determined based on the effects of the engine weight and/or the thrust provided by the engine (10) at the, or each, operating point.
6. The method (100, 200) of any one of the preceding claims, wherein the operation casing profile (50) is determined based on rapid transient effects.and optionally wherein the rapid transient effects include wind gust loading.
7. The method (100, 200) of any one of the preceding claims, further comprising repeating at least steps b), d) and e) for a plurality of iterations, wherein the resulting modified casing profile (56) for each iteration is used as the initial casing profile (46) for the following iteration, and optionally wherein the method further comprises: repeating at least steps b), d) and e) until a constant modified casing profile (56) is determined.
8. The method (100, 200) of any one of the preceding claims, wherein the casing profile defines the shape of the internal surface (8) of the housing (21) as the radial distance of the inner surface (8) from a rotational axis (9) of the engine (10) at associated angles of rotation around the rotational axis (9).
9. The method (100, 200) of Claim 8, wherein: the shape of the internal surface of the housing (21) is defined by the casing profiles at a plurality of axial points at or between the axial positions of the leading edges and the trailing edges of the fan blade tips; and the operating fan blade tip profile defines corresponding radii of the tips of the fan blades at or between the axial positions of the leading and trailing edges of the fan blade tips.
10. The method (100, 200) of any one of the preceding claims, wherein the modified casing profile (56) is non-axisym metric about a rotational axis of the engine.
11. The method (100, 200) of Claim 10, wherein the modified casing profile has a minor axis and a major axis, the minor axis and major axis being perpendicular to one another and the major axis being greater in length than the minor axis, and optionally wherein: i) the gas turbine engine has a vertical axis, and the minor axis of the modified casing profile (56) is orientated in a 90-degree range measured clockwise from the vertical axis of the engine when viewed in a rearward direction from a position forward of the fan; and/or ii) the modified casing profile (56) has a generally elliptical shape.
12. The method (100, 200) of Claim 10 or Claim 11, wherein the modified casing profile varies in radius from axisymmetric by 15 mm or less at any point around the rotational axis of the engine.
13. The method (100, 200) of any one of the preceding claims, wherein: a) the fan (23) comprises a plurality of fan blades (41), wherein each of the fan blades (41) is formed at least in part from a composite material, and optionally wherein each fan blade (41) comprises a body formed from a composite material and a leading edge formed from a metallic material; and/or b) the fan (23) has a diameter in the range of between 240 cm and 380 cm, preferably between 240 cm and 280 cm or between 330 cm and 380 cm.
14. The method (100, 200) of any one of the preceding claims, wherein: the gas turbine engine comprises an engine core (11) comprising a turbine (19), a compressor (14), and a core shaft (26) connecting the turbine to the compressor; the fan (23) is located upstream of the engine core; and the gas turbine engine further comprises a gearbox (30) that receives an input from the core shaft (26) and outputs drive to the fan (23) so as to drive the fan at a lower rotational speed than the core shaft.
15. A method (400) of manufacturing a gas turbine engine (10) for an aircraft, the gas turbine engine comprising a fan (23) having a plurality of fan blades (41), and a housing surrounding the fan (23), the housing having an internal surface (8) facing the blades of the fan (23), the method comprising: obtaining (402) a modified casing profile (56) associated with the gas turbine engine (10), the modified casing profile being determined using the method of any preceding claim; and manufacturing (404) the housing for the fan (23) of the gas turbine engine (10), 113 the housing having an internal wall shaped according to the modified casing profile (56).
16. A gas turbine engine (10) for an aircraft, the gas turbine engine manufactured according to the method (400) of Claim 15.
17. A computer program having instructions adapted to carry out the method according to any one of Claims 1 to 14.
18. A computer readable medium, having a computer program recorded thereon, 20 wherein the computer program is adapted to make the computer execute the method of any one of Claims 1 to 14.
19. A gas turbine engine (10) for an aircraft, the gas turbine engine comprising: a fan (23) having a plurality of fan blades (41); and a housing surrounding the fan (23), the housing having an internal surface (8) facing the blades of the fan (23), wherein the internal surface (8) of the housing has a non-axisymmetric profile about a rotational axis of the engine.
20. The gas turbine engine of Claim 15, wherein: a) the non-axisymmetric profile has a minor axis and a major axis, the minor axis and major axis being perpendicular to one another and the major axis being greater in length than the minor axis, and optionally wherein the gas turbine engine (10) has a vertical axis, and the minor axis of the modified casing profile (56) is orientated in a 90 degree range measured clockwise from the vertical axis of the engine when viewed in a rearward direction from a position forward of the fan (23); b) the non-axisymmetric profile is generally elliptical in shape; b) the non-axisymmetric profile varies in radius from axisymmetric by 15 mm or less at any point around the rotational axis of the engine.